Embedded electric machine

ABSTRACT

A gas turbine engine includes a compressor section and a turbine section together defining a core air flowpath. Additionally, a rotary component is rotatable with at least a portion of the compressor section and at least a portion of the turbine section. An electric machine is mounted coaxially with the rotary component and positioned at least partially inward of the core air flowpath along a radial direction of the gas turbine engine. A cavity wall defines at least in part a buffer cavity surrounding at least a portion of the electric machine to thermally insulate the electric machine, e.g., from the relatively high temperatures within the core air flowpath.

PRIORITY INFORMATION

The present application claims priority to, as a continuation of, U.S.patent application Ser. No. 15/242,789 filed on Aug. 22, 2016, which isincorporated by reference herein.

FIELD OF THE INVENTION

The present subject matter relates generally to a gas turbine enginehaving an embedded electric machine, and to a propulsion system for anaeronautical device including the same.

BACKGROUND OF THE INVENTION

Typical aircraft propulsion systems include one or more gas turbineengines. For certain propulsion systems, the gas turbine enginesgenerally include a fan and a core arranged in flow communication withone another. Additionally, the core of the gas turbine engine generalincludes, in serial flow order, a compressor section, a combustionsection, a turbine section, and an exhaust section. In operation, air isprovided from the fan to an inlet of the compressor section where one ormore axial compressors progressively compress the air until it reachesthe combustion section. Fuel is mixed with the compressed air and burnedwithin the combustion section to provide combustion gases. Thecombustion gases are routed from the combustion section to the turbinesection. The flow of combustion gasses through the turbine sectiondrives the turbine section and is then routed through the exhaustsection, e.g., to atmosphere.

For certain aircraft, it may be beneficial for the propulsion system toinclude an electric fan to supplement propulsive power provided by theone or more gas turbine engines included with the propulsion system.However, providing the aircraft with a sufficient amount of energystorage devices to power the electric fan may be space and weightprohibitive. Notably, certain gas turbine engines may include auxiliarygenerators positioned, e.g., within a cowling of the gas turbine engine.However, these auxiliary generators are not configured to provide asufficient amount of electrical power to adequately drive the electricfan.

Accordingly, a propulsion system for an aircraft having one or more gasturbine engines and electric generators capable of providing an electricfan, or other electric propulsor, with a desired amount of electricalpower would be useful.

BRIEF DESCRIPTION OF THE INVENTION

Aspects and advantages of the invention will be set forth in part in thefollowing description, or may be obvious from the description, or may belearned through practice of the invention.

In one exemplary embodiment of the present disclosure, a gas turbineengine is provided. The gas turbine engine defines a radial directionand an axial direction. The gas turbine engine includes a compressorsection and a turbine section arranged in serial flow order, thecompressor section and turbine section together defining a core airflowpath. The gas turbine engine also includes a rotary componentrotatable with at least a portion of the compressor section and with atleast a portion of the turbine section. The gas turbine engine alsoincludes an electric machine coupled to the rotary component at leastpartially inward of the core air flowpath along the radial direction.The gas turbine engine also includes a cavity wall defining at least inpart a buffer cavity, the buffer cavity surrounding at least a portionof the electric machine to thermally insulate the electric machine.

In another exemplary embodiment of the present disclosure, a propulsionsystem is provided for an aeronautical device. The propulsion systemincludes an electric propulsor and a gas turbine engine defining aradial direction and an axial direction. The gas turbine engine includesa compressor section and a turbine section arranged in serial floworder, the compressor section and turbine section together defining acore air flowpath. The gas turbine engine also includes a rotarycomponent rotatable with at least a portion of the compressor sectionand with at least a portion of the turbine section. The gas turbineengine also includes an electric machine coupled to the rotary componentat least partially inward of the core air flowpath along the radialdirection, the electric machine electrically connected to the electricpropulsor. The gas turbine engine also includes a cavity wall definingat least in part a buffer cavity, the buffer cavity surrounding at leasta portion of the electric machine to thermally insulate the electricmachine.

These and other features, aspects and advantages of the presentinvention will become better understood with reference to the followingdescription and appended claims. The accompanying drawings, which areincorporated in and constitute a part of this specification, illustrateembodiments of the invention and, together with the description, serveto explain the principles of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS

A full and enabling disclosure of the present invention, including thebest mode thereof, directed to one of ordinary skill in the art, is setforth in the specification, which makes reference to the appendedfigures, in which:

FIG. 1 is a top view of an aircraft according to various exemplaryembodiments of the present disclosure.

FIG. 2 is a port side view of the exemplary aircraft of FIG. 1

FIG. 3 is a schematic, cross-sectional view of a gas turbine engine inaccordance with an exemplary aspect of the present disclosure.

FIG. 4 is a schematic, cross-sectional view of an electric machineembedded in a gas turbine engine in accordance with an exemplaryembodiment of the present disclosure.

FIG. 5 is a schematic, cross-sectional view of an electric machineembedded in a gas turbine engine in accordance with another exemplaryembodiment of the present disclosure.

FIG. 6 is a close-up, cross-sectional view of an electric cablepositioned within a cooling conduit in accordance with an exemplaryembodiment of the present disclosure.

FIG. 7 is a schematic, cross-sectional view of an electric machineembedded in a gas turbine engine in accordance with yet anotherexemplary embodiment of the present disclosure.

FIG. 8 is a schematic, cross-sectional view of an electric machineembedded in a gas turbine engine in accordance with still anotherexemplary embodiment of the present disclosure.

FIG. 9 is a close-up, cross-sectional view of an electric cable inaccordance with an exemplary embodiment of the present disclosure.

FIG. 10 is a schematic, cross-sectional view of a gas turbine engine inaccordance with another exemplary embodiment of the present disclosure.

DETAILED DESCRIPTION OF THE INVENTION

Reference will now be made in detail to present embodiments of theinvention, one or more examples of which are illustrated in theaccompanying drawings. The detailed description uses numerical andletter designations to refer to features in the drawings. Like orsimilar designations in the drawings and description have been used torefer to like or similar parts of the invention. As used herein, theterms “first”, “second”, and “third” may be used interchangeably todistinguish one component from another and are not intended to signifylocation or importance of the individual components. The terms “forward”and “aft” refer to relative positions within a gas turbine engine, withforward referring to a position closer to an engine inlet and aftreferring to a position closer to an engine nozzle or exhaust. The terms“upstream” and “downstream” refer to the relative direction with respectto fluid flow in a fluid pathway. For example, “upstream” refers to thedirection from which the fluid flows, and “downstream” refers to thedirection to which the fluid flows.

The present application is directed generally towards a gas turbineengine of a propulsion system for an aircraft having an electric machineembedded therein. In at least certain embodiments, the gas turbineengine includes a compressor section and a turbine section arranged inserial flow order and together defining a core air flowpath. A rotarycomponent, such as a shaft or spool, is rotatable with at least aportion of the compressor section and the turbine section. The gasturbine engine additionally includes an electric machine embedded withinthe gas turbine engine. For example, the electric machine is rotatablewith the rotary component and is positioned coaxially with the rotarycomponent at least partially inward of the core air flowpath along aradial direction of the gas turbine engine. For example, in at leastcertain embodiments, the electric machine may be an electric generator,driven by the rotary component. Additionally, the gas turbine engineincludes a cavity wall defining at least in part a buffer cavity. Thebuffer cavity surrounds at least a portion of the electric machine tothermally insulate the electric machine from, e.g., relatively hightemperatures within the core air flowpath of the gas turbine engine.

Referring now to the drawings, wherein identical numerals indicate thesame elements throughout the figures, FIG. 1 provides a top view of anexemplary aircraft 10 as may incorporate various embodiments of thepresent invention. FIG. 2 provides a port side view of the aircraft 10as illustrated in FIG. 1. As shown in FIGS. 1 and 2 collectively, theaircraft 10 defines a longitudinal centerline 14 that extendstherethrough, a vertical direction V, a lateral direction L, a forwardend 16, and an aft end 18. Moreover, the aircraft 10 defines a mean line15 extending between the forward end 16 and aft end 18 of the aircraft10. As used herein, the “mean line” refers to a midpoint line extendingalong a length of the aircraft 10, not taking into account theappendages of the aircraft 10 (such as the wings 20 and stabilizersdiscussed below).

Moreover, the aircraft 10 includes a fuselage 12, extendinglongitudinally from the forward end 16 of the aircraft 10 towards theaft end 18 of the aircraft 10, and a pair of wings 20. As used herein,the term “fuselage” generally includes all of the body of the aircraft10, such as an empennage of the aircraft 10. The first of such wings 20extends laterally outwardly with respect to the longitudinal centerline14 from a port side 22 of the fuselage 12 and the second of such wings20 extends laterally outwardly with respect to the longitudinalcenterline 14 from a starboard side 24 of the fuselage 12. Each of thewings 20 for the exemplary embodiment depicted includes one or moreleading edge flaps 26 and one or more trailing edge flaps 28. Theaircraft 10 further includes a vertical stabilizer 30 having a rudderflap 32 for yaw control, and a pair of horizontal stabilizers 34, eachhaving an elevator flap 36 for pitch control. The fuselage 12additionally includes an outer surface or skin 38. It should beappreciated however, that in other exemplary embodiments of the presentdisclosure, the aircraft 10 may additionally or alternatively includeany other suitable configuration of stabilizer that may or may notextend directly along the vertical direction V or horizontal/lateraldirection L.

The exemplary aircraft 10 of FIGS. 1 and 2 includes a propulsion system100, herein referred to as “system 100”. The exemplary system 100includes one or more aircraft engines and one or more electricpropulsion engines. For example, the embodiment depicted includes aplurality of aircraft engines, each configured to be mounted to theaircraft 10, such as to one of the pair of wings 20, and an electricpropulsion engine. More specifically, for the embodiment depicted, theaircraft engines are configured as gas turbine engines, or rather asturbofan jet engines 102, 104 attached to and suspended beneath thewings 20 in an under-wing configuration. Additionally, the electricpropulsion engine is configured to be mounted at the aft end of theaircraft 10, and hence the electric propulsion engine depicted may bereferred to as an “aft engine.” Further, the electric propulsion enginedepicted is configured to ingest and consume air forming a boundarylayer over the fuselage 12 of the aircraft 10. Accordingly, theexemplary aft engine depicted may be referred to as a boundary layeringestion (BLI) fan 106. The BLI fan 106 is mounted to the aircraft 10at a location aft of the wings 20 and/or the jet engines 102, 104.Specifically, for the embodiment depicted, the BLI fan 106 is fixedlyconnected to the fuselage 12 at the aft end 18, such that the BLI fan106 is incorporated into or blended with a tail section at the aft end18, and such that the mean line 15 extends therethrough. It should beappreciated, however, that in other embodiments the electric propulsionengine may be configured in any other suitable manner, and may notnecessarily be configured as an aft fan or as a BLI fan.

Referring still to the embodiment of FIGS. 1 and 2, in certainembodiments the propulsion system further includes one or more electricgenerators 108 operable with the jet engines 102, 104. For example, oneor both of the jet engines 102, 104 may be configured to providemechanical power from a rotating shaft (such as an LP shaft or HP shaft)to the electric generators 108. Although depicted schematically outsidethe respective jet engines 102, 104, in certain embodiments, theelectric generators 108 may be positioned within a respective jet engine102, 104. Additionally, the electric generators 108 may be configured toconvert the mechanical power to electrical power. For the embodimentdepicted, the propulsion system 100 includes an electric generator 108for each jet engine 102, 104, and also includes a power conditioner 109and an energy storage device 110. The electric generators 108 may sendelectrical power to the power conditioner 109, which may transform theelectrical energy to a proper form and either store the energy in theenergy storage device 110 or send the electrical energy to the BLI fan106. For the embodiment depicted, the electric generators 108, powerconditioner 109, energy storage device 110, and BLI fan 106 are all areconnected to an electric communication bus 111, such that the electricgenerator 108 may be in electrical communication with the BLI fan 106and/or the energy storage device 110, and such that the electricgenerator 108 may provide electrical power to one or both of the energystorage device 110 or the BLI fan 106. Accordingly, in such anembodiment, the propulsion system 100 may be referred to as agas-electric propulsion system.

It should be appreciated, however, that the aircraft 10 and propulsionsystem 100 depicted in FIGS. 1 and 2 is provided by way of example onlyand that in other exemplary embodiments of the present disclosure, anyother suitable aircraft 10 may be provided having a propulsion system100 configured in any other suitable manner. For example, it should beappreciated that in various other embodiments, the BLI fan 106 mayalternatively be positioned at any suitable location proximate the aftend 18 of the aircraft 10. Further, in still other embodiments theelectric propulsion engine may not be positioned at the aft end of theaircraft 10, and thus may not be configured as an “aft engine.” Forexample, in other embodiments, the electric propulsion engine may beincorporated into the fuselage of the aircraft 10, and thus configuredas a “podded engine,” or pod-installation engine. Further, in stillother embodiments, the electric propulsion engine may be incorporatedinto a wing of the aircraft 10, and thus may be configured as a “blendedwing engine.” Moreover, in other embodiments, the electric propulsionengine may not be a boundary layer ingestion fan, and instead may bemounted at any suitable location on the aircraft 10 as a freestreaminjection fan. Furthermore, in still other embodiments, the propulsionsystem 100 may not include, e.g., the power conditioner 109 and/or theenergy storage device 110, and instead the generator(s) 108 may bedirectly connected to the BLI fan 106.

Referring now to FIG. 3, a schematic cross-sectional view of apropulsion engine in accordance with an exemplary embodiment of thepresent disclosure is provided. In certain exemplary embodiments, thepropulsion engine may be configured a high-bypass turbofan jet engine200, herein referred to as “turbofan 200.” Notably, in at least certainembodiments, the jet engines 102, 104 may be also configured ashigh-bypass turbofan jet engines. In various embodiments, the turbofan200 may be representative of jet engines 102, 104. Alternatively,however, in other embodiments, the turbofan 200 may be incorporated intoany other suitable aircraft 10 or propulsion system 100.

As shown in FIG. 3, the turbofan 200 defines an axial direction A(extending parallel to a longitudinal centerline 201 provided forreference), a radial direction R, and a circumferential direction C(extending about the axial direction A; not depicted in FIG. 3). Ingeneral, the turbofan 200 includes a fan section 202 and a core turbineengine 204 disposed downstream from the fan section 202.

The exemplary core turbine engine 204 depicted generally includes asubstantially tubular outer casing 206 that defines an annular inlet208. The outer casing 206 encases, in serial flow relationship, acompressor section including a booster or low pressure (LP) compressor210 and a high pressure (HP) compressor 212; a combustion section 214; aturbine section including a high pressure (HP) turbine 216 and a lowpressure (LP) turbine 218; and a jet exhaust nozzle section 220. Thecompressor section, combustion section 214, and turbine section togetherdefine a core air flowpath 221 extending from the annular inlet 208through the LP compressor 210, HP compressor 212, combustion section214, HP turbine section 216, LP turbine section 218 and jet nozzleexhaust section 220. A high pressure (HP) shaft or spool 222 drivinglyconnects the HP turbine 216 to the HP compressor 212. A low pressure(LP) shaft or spool 224 drivingly connects the LP turbine 218 to the LPcompressor 210.

For the embodiment depicted, the fan section 202 includes a variablepitch fan 226 having a plurality of fan blades 228 coupled to a disk 230in a spaced apart manner. As depicted, the fan blades 228 extendoutwardly from disk 230 generally along the radial direction R. Each fanblade 228 is rotatable relative to the disk 230 about a pitch axis P byvirtue of the fan blades 228 being operatively coupled to a suitableactuation member 232 configured to collectively vary the pitch of thefan blades 228 in unison. The fan blades 228, disk 230, and actuationmember 232 are together rotatable about the longitudinal axis 12 by LPshaft 224 across a power gear box 234. The power gear box 234 includes aplurality of gears for stepping down the rotational speed of the LPshaft 224 to a more efficient rotational fan speed.

Referring still to the exemplary embodiment of FIG. 3, the disk 230 iscovered by rotatable front hub 236 aerodynamically contoured to promotean airflow through the plurality of fan blades 228. Additionally, theexemplary fan section 202 includes an annular fan casing or outernacelle 238 that circumferentially surrounds the fan 226 and/or at leasta portion of the core turbine engine 204. The nacelle 238 is supportedrelative to the core turbine engine 204 by a plurality ofcircumferentially-spaced outlet guide vanes 240. A downstream section242 of the nacelle 238 extends over an outer portion of the core turbineengine 204 so as to define a bypass airflow passage 244 therebetween.

Although not depicted, the variety of rotatory components of theturbofan engine 10 (e.g., LP shaft 224, HP shaft 222, fan 202) may besupported by one or more oil lubricated bearings. The turbofan engine 10depicted includes a lubrication system 245 for providing one or more ofthe oil lubricated bearings with lubrication oil. Further, thelubrication system 245 may include one or more heat exchangers fortransferring heat from the lubrication oil with, e.g., bypass air, bleedair, or fuel.

Additionally, the exemplary turbofan 200 depicted includes an electricmachine 246 rotatable with the fan 226. Specifically, for the embodimentdepicted, the electric machine 246 is configured as an electricgenerator co-axially mounted to and rotatable with the LP shaft 224 (theLP shaft 224 also rotating the fan 226 through, for the embodimentdepicted, the power gearbox 234). As used herein, “co-axially” refers tothe axes being aligned. It should be appreciated, however, that in otherembodiments, an axis of the electric machine 246 may be offset radiallyfrom the axis of the LP shaft 224 and further may be oblique to the axisof the LP shaft 224, such that the electric machine 246 may bepositioned at any suitable location at least partially inward of thecore air flowpath 221.

The electric machine 246 includes a rotor 248 and a stator 250. Incertain exemplary embodiments, the rotor 248 and stator 250 of theelectric machine 246 are configured in substantially the same manner asthe exemplary rotor and stator of the electric machine described below.Notably, when the turbofan engine 200 is integrated into the propulsionsystem 100 described above with reference to FIGS. 1 and 2, the electricgenerators 108 may be configured in substantially the same manner as theelectric machine 246 of FIG. 3.

It should be also appreciated, however, that the exemplary turbofanengine 200 depicted in FIG. 3 is provided by way of example only, andthat in other exemplary embodiments, the turbofan engine 200 may haveany other suitable configuration. For example, in other exemplaryembodiments, the turbofan engine 200 may be configured as a turbopropengine, a turbojet engine, a differently configured turbofan engine, orany other suitable gas turbine engine.

Referring now to FIG. 4, an electric machine 246 embedded within a gasturbine engine in accordance with an exemplary embodiment of the presentdisclosure is depicted. More particularly, for the embodiment depicted,the electric machine 246 is embedded within a turbine section of the gasturbine engine, and more particularly still, is attached to an LP shaft224 of the gas turbine engine. Additionally, the electric machine 246 ispositioned at least partially within or aft of the turbine section alongan axial direction A. In certain exemplary embodiments, the electricmachine 246 and gas turbine engine depicted in FIG. 4 may be configuredin substantially the same manner as the exemplary electric machine 246and turbofan engine 200 described above with reference to FIG. 3.Accordingly, the same or similar numbers may refer to the same orsimilar parts.

As is depicted, the electric machine 246 generally includes a rotor 248and a stator 250. The rotor 248 is attached via a plurality of rotorconnection members 252 directly to the LP shaft 224, such that the rotor248 is rotatable with the LP shaft 224. By contrast, the stator 250 isattached via one or more stator connection members 254 to a structuralsupport member 256 of the turbine section. In at least certain exemplaryembodiments, the electric machine 246 may be an electric generator, suchthat the rotor 248, and rotor connection members 252, are driven by theLP shaft 224. With such an embodiment, a rotation of the rotor 248relative to the stator 250 may generate electrical power, which may betransferred via an electric communication bus 258, discussed in greaterdetail below.

It should be appreciated, however, that in other exemplary embodiments,the electric machine 246 may instead have any other suitableconfiguration. For example, in other embodiments the electric machine246 may include the rotor 248 located radially inward of the stator 250(e.g., as an in-running electric machine).

Referring still to the exemplary electric machine 246 of FIG. 4, thestructural support member 256 may be configured as part of an aft frameassembly 257 and extends from an aft frame strut 258 of the aft frameassembly 257 of the gas turbine engine. The aft strut 258 extendsthrough the core air flowpath 221 of the gas turbine engine, and isconfigured to provide structural support for the gas turbine engine. Thestructural support member 256 also extends forward to support an aftengine bearing 262—the aft engine bearing 262 rotatably supporting anaft end of the LP shaft 224.

The stator connection member 254 may be an annular/cylindrical memberextending from the structural support member 256 of the gas turbineengine. For the embodiment depicted, the stator connection member 254supports rotation of the plurality of rotor connection members 252through one or more bearings. More specifically, a forward electricmachine bearing 264 is positioned forward of the electric machine 246and between the rotor connection member 252 and the stator connectionmember 254 along a radial direction R. Similarly, an aft electricmachine bearing 266 is positioned aft of the electric machine 246 andbetween the rotor connection member 252 and the stator connection member254 along the radial direction R. Particularly for the embodimentdepicted, the forward electric machine bearing 264 is configured as aroller element bearing and the aft electric machine bearing 266 includesa pair of bearings, the pair of bearings configured as a roller elementbearing and a ball bearing. It should be appreciated, however, that theforward and aft electric machine bearings 264, 266 may in otherembodiments, have any other suitable configuration and the presentdisclosure is not intended to be limited to the specific configurationdepicted, unless such limitations are added to the claims.

The gas turbine engine further includes a cavity wall 268 surrounding atleast a portion of the electric machine 246. More specifically, for theembodiment depicted, the cavity wall 268 substantially completelysurrounds electric machine 246, extending from a location forward of theelectric machine 246 (attached to the structural support member 256,through the stator connection member 254) to a location aft of theelectric machine 246. The cavity wall 268 defines at least in part anelectric machine sump 270 substantially completely surrounding theelectric machine 246. More specifically, the electric machine sump 270extends from a location forward of the electric machine 246 continuouslyto a location aft of the electric machine 246. Certain components of thegas turbine engine include openings 272 to allow for such a continuousextension of the electric machine sump 270.

Notably, for the embodiment depicted, the electric machine sump 270additionally encloses the aft engine bearing 262 of the gas turbineengine. The gas turbine engine includes a sealing arm 274 attached tothe structural support member 256 and extending forward of the aftengine bearing 262 to form a seal with the LP shaft 224 and include theaft engine bearing 262 within the electric machine sump 270. Notably, aseal assembly 276 is provided as part of the sealing arm 274 and/or theLP shaft 224 for providing such a seal and maintaining a sealed electricmachine sump 270. As is also depicted, the gas turbine engine furtherincludes a plurality of seal assemblies 276 adjacent to the forwardelectric machine bearing 264 and the aft electric machine bearings 266,for maintaining a sealed electric machine 246, i.e., preventinglubrication oil from reaching the rotor 248 and stator 250 of theelectric machine 246.

Moreover, the gas turbine engine depicted includes an electric machinelubrication system 278, with the electric machine lubrication system 278in fluid communication with the electric machine sump 270, for providinga thermal fluid to the electric machine sump 270. For the embodimentdepicted, the electric machine lubrication system 278 may operateindependently of a gas turbine engine lubrication system, such as thelubrication system 245 described above with reference to FIG. 3.

Specifically, for the embodiment depicted, the electric machinelubrication system 278 include a supply pump 280 connected to a supplyline 282 extending to the electric machine sump 270. The supply line 282extends from a location outward of the core air flowpath 221 along theradial direction R, through the aft engine strut 258 (and through thecore air flowpath 221), through the cavity wall 268 and to the electricmachine sump 270. The thermal fluid may be a lubrication oil or othersuitable lubricant for lubricating the forward electric machine bearing264 and the aft electric machine bearings 266, as well as the aft enginebearing 262. Notably, the thermal fluid is further configured to acceptheat from the plurality of bearings and the electric machine sump 270.The heated thermal fluid is scavenged out of the electric machine sump270 via a scavenge line 284 of the lubrication system 278, the scavengeline 284 extending from the electric machine sump 270, through the coreair flowpath 221, and to a scavenge pump 286. It should be appreciated,however, that although the scavenge line 284 is, for the embodimentdepicted, extending through the core air flowpath 221 at a locationoutside of the strut 260, in other embodiments, the scavenge line 284may instead extend through the strut 260 alongside the supply line 282.

Notably, for the embodiment depicted, the electric machine lubricationsystem 278, including the supply pump 280 and scavenge pump 286, may bepowered at least in part by the electric machine 246. Additionally,although not depicted, the electric machine lubrication system 278 mayfurther include one or more heat exchangers for reducing a temperatureof the scavenged thermal fluid, before such thermal fluid is providedback through the supply line 282 to the electric machine sump 270.

Notably, with such an embodiment, the lubrication system 278 may furtherbe configured as part of a cooling system of the gas turbine engine forreducing a temperature of the electric machine 246. For example, theinventors of the present disclosure have discovered that for at leastcertain embodiments, providing lubrication oil to the lubrication oilsupply line 282 at a temperature less than about 275° F., such as lessthan about 250° F., may allow for the lubrication oil to accept anamount of heat necessary to maintain the electric machine 246 within adesired temperature operating range during operation of the gas turbineengine. It should be appreciated, that as used herein, terms ofapproximation, such as “about” or “approximately,” refer to being withina 10% margin of error. Also, it should be appreciated, that in otherembodiments, the lubrication oil provided to the supply line 282 mayhave any other suitable temperature.

In order to further maintain a temperature of the electric machine 246,the cooling system of exemplary gas turbine engine depicted furtherincludes a buffer cavity 288 surrounding at least a portion of theelectric machine 246 to thermally insulate the electric machine 246.More specifically, for the embodiment depicted, the cavity wall 268 alsoat least partially defines the buffer cavity 288 with the buffer cavity288 being positioned opposite the cavity wall 268 of the electricmachine sump 270. Additionally, as is depicted in FIG. 4, an extensionmember 290 is attached to or formed integrally with the structuralsupport member 256 and extends at least partially around the cavity wall268. Specifically, for the embodiment depicted, the structural supportmember 256 and extension member 290 together extend completely aroundthe cavity wall 268. The structural support member 256 and extensionmember 290 together define the buffer cavity 288, which for theembodiment depicted extends continuously from a location forward of theelectric machine 246 to a location aft of the electric machine 246 alongthe axial direction A. The buffer cavity 288 may act as an insulatorfrom relatively hot operating temperatures within the core air flowpath221 extending through the turbine section of the gas turbine engine.

Furthermore, for the embodiment depicted, the gas turbine engine furtherincludes a cooling duct 292. The cooling duct 292 is in airflowcommunication with the buffer cavity 288 for providing a cooling airflowto the buffer cavity 288. For example, in the embodiment depicted, thecooling duct 292 defines an outlet 293 extending through the structuralsupport member 256 for providing the cooling airflow from the coolingduct 292 through the structural support member 256 and into the buffercavity 288. The cooling duct 292 may also be in airflow communicationwith a relatively cool air source for providing the cooling airflow. Incertain exemplary embodiments, the cool air source may be a compressorsection of the gas turbine engine (wherein the cooling airflow may bediverted from the compressor section), or a fan of the gas turbineengine (wherein the cooling airflow may be diverted from the fan).Notably, for the embodiment depicted, the gas turbine engine furtherincludes an exhaust duct 291. The exhaust duct 291 is in airflowcommunication with the buffer cavity 288 and is configured to exhaustthe cooling airflow to the core air flowpath 221, a bypass passage(e.g., passage 244 of FIG. 3), or an ambient location. Such aconfiguration may allow for a continuous cooling airflow through thebuffer cavity 288.

As discussed, the electric machine lubrication system 278, cooling duct292, and buffer cavity 288 are each configured as part of the coolingsystem for maintaining at least certain components of the electricmachine 246 within a desired temperature range. For example, for theembodiments wherein the electric machine 246 is configured as anelectric generator, the electric generator may be configured as apermanent magnet electric generator including a plurality of permanentmagnets 294 (depicted in phantom). For these embodiments, the rotor 248may include the plurality of permanent magnets 294 and the stator 250may include one or more coils of electrically conductive wire (notshown). It should be appreciated, however, that in other embodiments,the electric machine 246 may alternatively be configured as anelectromagnetic generator, including a plurality of electromagnets andactive circuitry, as an induction type electric machine, a switchedreluctance type electric machine, as a synchronous AC electric machine,or as any other suitable electric generator or motor.

As will be appreciated, each of the plurality of permanent magnets 294,when included, defines a Curie temperature limit, which may be less thana temperature within the core air flowpath 221 extending through theturbine section of the gas turbine engine. The cooling system of the gasturbine engine maintains a temperature of the electric machine 246, andmore particularly each of the permanent magnets 294, below the Curietemperature limit for the plurality of permanent magnets 294. Further,the cooling system may maintain a temperature of the electric machine246 below a predetermined limit of the Curie temperature limit to, e.g.,increase a useful life of the electric machine 246. For example, incertain exemplary embodiments, the cooling system the gas turbine enginemay maintain a temperature of the electric machine 246 below at leastabout a 50 degrees Fahrenheit (° F.) limit of the Curie temperaturelimit, such as below at least about a 75° F. limit or 100° F. limit ofthe Curie temperature limit. Maintaining a temperature of the electricmachine 246 below such a limit of the Curie temperature limit mayfurther prevent any permanent magnets of the electric machine 246 fromexperiencing un-recoverable (or permanent) de-magnetization, which mayhave a negative life impact on the electric machine 246.

It should be appreciated, however, that the exemplary cooling systemdepicted in the embodiment of FIG. 4 is provided by way of example only.In other embodiments, the gas turbine engine may include any othersuitable cooling system. For example, in other embodiments, the electricmachine lubrication system 278 may have any other suitableconfiguration. For example, the electric machine lubrication system 278may be operable with the engine lubrication system 278. Additionally, incertain embodiments, the cavity wall 268 may have any other suitablefeatures for maintaining a temperature of the electric machine 246within a desired operating range. For example, referring now briefly toFIG. 5, a cross-sectional, schematic view of an electric machine 246embedded within a gas turbine engine in accordance with anotherexemplary embodiment of the present disclosure is depicted. Theexemplary gas turbine engine depicted in FIG. 5 may be configured insubstantially the same manner as the exemplary gas turbine enginedepicted in FIG. 4, and accordingly the same or similar numbers mayrefer to same or similar part. However, for the embodiment of FIG. 5,the cavity wall 268, which at least partially defines a buffer cavity288, further includes a layer 296 of insulation to further insulate theelectric machine 246 from relatively hot operating temperatures withinthe core air flowpath 221 extending through the turbine section of thegas turbine engine. The insulation layer 296 may be any suitableinsulation for reducing a thermal conductivity of the cavity wall 268surrounding the electric machine 246. Additionally, although notdepicted, in certain embodiments, a portion of the structural supportmember 256 and extension member 290 (also at least partially definingthe buffer cavity 288) may also include a layer of insulation.

Referring again to the embodiment of FIG. 4, as briefly discussed aboveduring operation of the gas turbine engine, the LP shaft 224 may rotatethe rotor 248 of the electric machine 246, allowing electric machine 246to function as an electric generator producing electrical power.Additionally, the electric machine 246 is in electrical communicationwith—i.e. electrically connected to—the electric communication bus 258.The electric communication bus 258 is electrically connected to theelectric machine 246 at a location radially inward of the core airflowpath 221. The electric communication bus 258 includes a firstjuncture box 298 mounted to the stator connection member 254. The firstjuncture box 298 receives an electrical line 300 from the electricmachine 246 (for the embodiment depicted, from the stator 250 of theelectric machine 246) and connects the electric line 300 to anintermediate section 302 of the electric communication bus 258. Theintermediate section 302 extends through the core air flowpath 221 to asecond juncture box 304 mounted at a location radially outward of thecore air flowpath 221, within a cowling of the gas turbine engine. Thesecond juncture box 304 connects the intermediate section 302 of theelectric communication bus 258 to an outlet line 306 of the electriccommunication bus 258 for connection to one or more systems of the gasturbine engine and/or aircraft with which the gas turbine engine isinstalled. As briefly mentioned above, the electric machine lubricationsystem 278 may be electrically connected to the outlet line 306 of theelectric communication bus 258 for powering the electric machinelubrication system 278.

As stated and depicted in FIG. 4, at least a portion of the electriccommunication bus 258 extends through the core air flowpath 221. Morespecifically, for the embodiment depicted, the intermediate section 302of the electric communication bus 258 extends through the core airflowpath 221 at a location downstream of a combustion section of the gasturbine engine (such as the combustion section 214 of the exemplaryturbofan engine 200 of FIG. 3). In particular, the intermediate section302 extends through/is positioned within the aft strut 258—the aft strut258 located in a portion of the core air flowpath 221 immediatelydownstream of the HP turbine 216.

Moreover, as is depicted schematically, the exemplary intermediatesection 302 depicted is a cooled portion of the electric communicationbus 258, including an electric cable 308 (i.e., an electric conductor)positioned within/extending through a conduit containing a coolingfluid. Specifically, reference will now also be made to FIG. 6,providing a close-up view of a portion of the intermediate section 302that is configured to extend through the core air flowpath 221 of thegas turbine engine. As is depicted, the intermediate section 302 of theelectric communication bus 258 includes the electric cable 308positioned within and extending coaxially with the supply line 282, suchthat during operation, the electric cable 308 is surrounded byrelatively cool flow of thermal fluid (represented by arrows 310) to beprovided, e.g., to the electric machine sump 270. Accordingly, thesupply line 282 is considered for the embodiment depicted as part of theelectric machine lubrication system 278 as well as part of theintermediate section 302 of the electric communication bus 258. Duringoperation, the thermal fluid surrounding the electric cable 308 withinthe intermediate section 302 of the electric communication bus 258 mayprotect the electric cable 308 from relatively high temperatures withinthe core air flowpath 221, maintaining a temperature of the electriccable 308 within a desired operating range. It should be appreciated,however, that in other embodiments, the intermediate section 302 of theelectric communication bus 258 may instead include the electric cable308 positioned within and extending coaxially with the scavenge line 284(which may also extend through the strut 260 in certain embodiments).

Notably, the electric cable 308 may be any suitable cable 308, and forthe embodiment depicted includes an electrical insulation layer 312surrounding a conducting core portion 314. The electrical insulationlayer 312 may include any suitable electrical insulation capable ofbeing exposed to the relatively high temperatures and further capable ofinsulating relatively high amounts of electrical power which may betransported through the conducting core portion 314 of the electriccable 308 (see discussion below). Additionally, although not depicted,the electric cable 308 may additionally include a barrier layersurrounding the electric insulation layer 312 and conducting coreportion 314 to prevent lubrication oil from contacting the insulationlayer 312 and conducting core portion 314. Additionally, still, incertain embodiments, the electric cable 308 may be configured insubstantially the same manner as the electric cable 308 described belowwith reference to FIG. 9.

As will be discussed in greater detail below, the intermediate section302 of the electric communication bus 258 is configured to transferrelatively high power levels of electrical power. Accordingly, duringoperation, the intermediate section 302 of the electric communicationbus 258 may experience a relatively high amount of Joule heating, orresistive heating, as a result of the relatively high power levels beingtransferred. Positioning the electric cable 308 of the intermediatesection 302 coaxially with the lubrication oil supply line 282 mayassist with maintaining a temperature of the electric cable 308 within adesired operating temperature range, despite the resistive heatingexperienced and exposure to the core air flowpath 221.

It should be appreciated, however, that in other exemplary embodiments,the electric communication bus 258 may have any other suitableconfiguration for transferring electrical power from the electricmachine 246 located radially inward from the core air flowpath 221 to alocation radially outward of the core air flowpath 221. For example,referring now briefly to FIG. 7, a cross-sectional, schematic view of anelectric machine 246 embedded within a gas turbine engine in accordancewith yet another exemplary embodiment of the present disclosure isdepicted. The exemplary gas turbine engine depicted in FIG. 7 may beconfigured in substantially the same manner as exemplary gas turbineengine depicted in FIG. 4, and accordingly the same or similar numbersmay refer to same or similar part.

However, for the embodiment of FIG. 7, the electric communication bus258 is instead configured as a superconducting, or hyper conducting,electric communication bus 258. Accordingly, for the embodiment of FIG.7, the intermediate section 302 of the electric communication bus 258may not be configured with the supply line 282 of the electric machinelubrication system 278. Instead, the exemplary electric communicationbus 258 includes a separate cooled conduit 316 within which the electriccable 308 is positioned and extends. The electric communication bus 258includes a refrigerant system 318 for providing a cold refrigerantwithin the cooled conduit 316 to maintain a temperature of the electriccable 308 extending therethrough at a relatively low temperature. Forexample, in certain embodiments, the refrigerant system may maintain atemperature of the electric cable 308 at or below a critical temperatureof the material forming the electric cable 308, or at least 1° F. coolerthan the critical temperature of the material forming the electric cable308.

Additionally, the cold refrigerant extends to a first juncture box 298,where the refrigerant is separated from the electric line in returnedthrough a return line 320 (partially depicted). For the embodimentdepicted, the electric communication bus 258 may additionally includecomponents for operating the refrigeration system 318 in a refrigerationcycle, such as a pump, a condenser, and an expansion valve (notdepicted). Notably, in at least certain embodiments, the portion of theintermediate section 302 extending through the core air flowpath 221 mayact as an evaporator of the refrigerant cycle.

Although for the embodiment depicted the gas turbine engine includes aseparate electric machine lubrication system 278 and refrigerant system318, in other embodiments the refrigerant utilized by the refrigerantsystem 318 of the electric communication bus 258 may additionally act asa lubricant for the various bearings within the electric machine 246(and for the embodiment depicted, for the aft engine bearing 262), suchthat the refrigerant system 318 and electric machine lubrication system278 may be configured together as a single system.

Referring now to FIG. 8, a cross-sectional, schematic view of anelectric machine 246 embedded within a gas turbine engine in accordancewith still another exemplary embodiment of the present disclosure isdepicted. The exemplary gas turbine engine depicted in FIG. 8 may beconfigured in substantially the same manner as exemplary gas turbineengine depicted in FIG. 4, and accordingly the same or similar numbersmay refer to same or similar part. However, for the embodiment of FIG.8, an intermediate section 302 of an electric communication bus 258 isnot configured coaxially with a cooling fluid conduit (e.g., a supplyline 282). Instead, for the embodiment of FIG. 8, the intermediatesection 302 of the electric communication bus 258 is formed of anelectric cable 308 designed to withstand the relatively hightemperatures of a core air flowpath 221 of the gas turbine engine at alocation downstream of a combustion section of the gas turbine engine.

More specifically, as with the embodiments described above, the electriccommunication bus 258 includes a first juncture box 298, a secondjuncture box 304, and the electric cable 308 extending therebetween(i.e., the intermediate section 302). Although the electric cable 308 isdepicted as a single cable, in certain embodiments, the electric cablemay include a plurality of electric cables. Referring now also brieflyto FIG. 9, providing a close-up, schematic view of the electric cable308, the electric cable 308 is formed of a material capable oftransmitting relatively high amounts of electrical power and beingexposed to the relatively high temperatures within the core air flowpath221 without oxidizing.

For example, in certain embodiments, the electric cable 308 may consistof at least one solid nickel wire core. Or, as in the embodimentdepicted, the cable 308 may consist of at least one high conductivitycore volume, such as a low resistivity/high conductivity cable core 322,and at least one dielectric (electrically-insulating) barrier volume,such as a high resistivity cable jacket 324. The cable core 322 ispositioned within the cable jacket 324, such that the cable jacket 324encloses the cable core 322. In certain exemplary embodiments, the cablecore 322 may be a copper core volume and the cable jacket 324 may be anon-copper jacket volume. The cable jacket 324 may be established by oneor more encasement processes, such as dipping, co-extrusion, plating,spraying, cladding, swaging, roll-forming, welding, or a combinationthereof. The electric cable 308 depicted additionally includes anoxidation barrier volume 323 positioned between the cable core 322 andcable jacket 324. Notably, the cable 308 may be configured as a wirebraid, a transposed and compacted wire bundle, transposed bundle(s) oftransposed wire bundle(s), or any other suitable cable configuration fortransferring alternating current (“AC”) power in a manner to reduce ACcoupling losses.

Additionally, for the embodiment depicted, the cable core 322 and cablejacket 324 of the electric cable 308 are covered and enclosed within ahigh temperature electric insulation material 326. For example, incertain embodiments, the high temperature electric insulation material326 may be a sprayed lamellar barrier coating (ceramic), at least onefractionally-overlapped tape layer (mica, glass fiber, ceramic fiber,and/or polymeric film), external armor barrier (braided, metallic and/ornon-metallic), or combinations thereof. The high temperature electricinsulation material 326 may be suitable for insulating cablestransferring relatively high amounts of electrical power at relativelyhigh temperatures, as discussed below. Further, for the embodimentdepicted, the electric cable 308 includes at least one external armorvolume 325 as an anti-abrasion barrier, which in certain embodiments maybe the same as the insulation material 326.

As is also depicted, the electric machine lubrication system 278(configured as part of the overall electric machine cooling system) isconfigured to provide thermal fluid directly to the second juncture box304 through a connection line 328 for actively cooling the secondjuncture box 304. Additionally, the thermal fluid supply line 282 of theelectric machine lubrication system 278 extends to the first juncturebox 298 and provides a flow of thermal fluid directly to the firstjuncture box 298 for actively cooling the first juncture box 298.Notably, for the embodiment depicted, the first juncture box 298includes a thermal fluid outlet 330 for ejecting the flow of thermalfluid provided thereto to the electric machine sump 270.

By actively cooling the first juncture box 298 and the second juncturebox 304, the intermediate section 302 including the electric cable 308may be allowed to operate at relatively high temperatures, such astemperatures resulting from exposure to the core air flowpath 221, aswell as from Joule heating, or electric resistance heating, of theelectric cable 308 during operation of the electric machine 246. Atemperature of the electric cable 308 with such a configuration may bereduced at the first juncture box 298 and at the second juncture box304, allowing for the electric cable 308 to be electrically connected toother electrical lines (e.g., outlet line 306 and electric line 300),which may not be configured for operating at the relatively hightemperatures at which the electric cable 308 of the intermediate section302 is capable of operating.

Moreover, as is also depicted, schematically, further beneficial coolingmay be achieved by equipping the second juncture box 304 with anembedded auxiliary fluid flow circuit 331 in heat transfer communicationwith the fluid transiting connection line 328. The auxiliary fluidwithin the auxiliary fluid flow circuit 331 may be the same fluidsupplied by the fluid supply line 282, or alternatively, may be adistinct thermal transfer fluid. Further, although not depicted, theauxiliary fluid may itself be in subsequent heat transfer communicationwith a heat-sinking media such as aircraft engine fuel, propulsor fanair, or a motor electronics coolant.

During operation of a gas turbine engine including an electric machine246 in accordance with an exemplary embodiment of the presentdisclosure, the electric machine 246 may be configured to generate arelatively high amount of alternating current electric power. Forexample, in certain embodiments, the electric machine 246 may beconfigured to generate and deliver through the electric communicationbus 258 electrical power at five hundred (500) Volts (“V”) or more. Forexample, in certain embodiments, the electric machine 246 may beconfigured to generate and deliver through the electric communicationbus 258 electrical power at six hundred (600) V or more. Such aconfiguration may be enabled by the disclosed cooling systems formaintaining a temperature of the electric machine 246 within a certainoperating temperature range, and/or by designing the intermediatesection 302 of the electric communication bus 258 in a manner allowingit to be exposed to the relatively high temperatures within the core airflowpath 221 downstream of the combustion section of the gas turbineengine.

Referring again briefly to FIGS. 1 and 2, in certain exemplaryembodiments of the present disclosure a propulsion system 100 isprovided having a plurality of gas turbine engines and electricmachines. For example, the propulsion system 100 may include a firstengine 102 and electric machine 108 and a second engine 104 and electricmachine 108. Each of the first and second engines 102, 104 andrespective electric machines 108 may be configured in substantially thesame manner as one or more of the gas turbine engines and embeddedelectric machines 246 described above with reference to FIGS. 4 through8. With such an exemplary embodiment, the first engine 102 and electricmachine 108 may be configured to generate electrical power at a firstvoltage level and the second engine 104 and electric machine 108 may beconfigured to generate electrical power at a second voltage level. Thefirst and second voltage levels generated may be provided through anelectric communication bus 111 to an electric propulsion device, such asthe exemplary BLI fan 106 depicted. Notably, in at least certainembodiments, the electric propulsion device may require (or desire)electrical power at a voltage level greater than each of the first andsecond engines 102, 104 and respective electric machines 108 may safelygenerate individually. Accordingly, in certain exemplary aspects, thefirst voltage level with respect to a ground plane of the aircraft 10may be a positive voltage level and the second voltage level may be at anegative voltage level with respect to the ground plane of the aircraft10. Further, in at least certain embodiments, the first voltage levelmay have substantially the same absolute value as an absolute value ofthe second voltage level. With such a configuration, the pair of firstand second engines 102, 104 and respective electric machines 108 maytherefore be capable of providing a net differential voltage toelectrical terminations of the electric propulsion device approximatelytwice as great as a single engine and electric machine may otherwise becapable of, therefore providing the electric propulsion device a desiredamount of electrical power.

Moreover, referring now to FIG. 10, a schematic, cross-sectional view isprovided of a gas turbine engine in accordance with another exemplaryembodiment of the present disclosure. In certain embodiments, theexemplary gas turbine engine depicted in FIG. 10 may be configured insubstantially the same manner as exemplary gas turbine engine describedabove with reference FIG. 3. Accordingly, the same or similar numbersmay refer to the same or similar part. For example, as is depicted, thegas turbine engine is configured as a turbofan engine generallycomprising a fan 202 and a core turbine engine 204. The core turbineengine 204 includes an LP compressor 210 connected to an LP turbine 218through an LP shaft 224, as well as an HP compressor 212 connected to anHP turbine 216 through an HP shaft 222. For the embodiment depicted, theturbofan engine 200 further includes an electric machine 246. Theelectric machine 246 may be configured in substantially the same manneras one or more of the embodiments described above with reference toFIGS. 4 through 9.

However, as is depicted schematically and in phantom, for the embodimentdepicted, the electric machine 246 may be positioned at any othersuitable location. For example, the electric machine 246 may be anelectric machine 246A coaxially mounted with the LP shaft 224 at alocation forward of the HP compressor 212 and substantially radiallyinward of the LP compressor 210. Additionally, or alternatively, theelectric machine 246 may be an electric machine 246B coaxially mountedwith the HP shaft 222, e.g., at a location forward of the HP compressor212. Additionally, or alternatively still, the electric machine 246 maybe an electric machine 246C coaxially mounted with the LP shaft 224 alocation at least partially aft of the HP turbine 216 and at leastpartially forward of the LP turbine 218. Additionally, or alternativelystill, the electric machine 246 may be an electric machine 246Dcoaxially mounted with the LP shaft 224 and the HP shaft 222, such thatthe electric machine 246D is a differential electric machine. Moreover,in still other embodiments, the electric machine 246 may be mounted atany other suitable location.

This written description uses examples to disclose the invention,including the best mode, and also to enable any person skilled in theart to practice the invention, including making and using any devices orsystems and performing any incorporated methods. The patentable scope ofthe invention is defined by the claims, and may include other examplesthat occur to those skilled in the art. Such other examples are intendedto be within the scope of the claims if they include structural elementsthat do not differ from the literal language of the claims, or if theyinclude equivalent structural elements with insubstantial differencesfrom the literal languages of the claims.

What is claimed is:
 1. A gas turbine engine defining a radial directionand an axial direction, the gas turbine engine comprising: a compressorsection and a turbine section arranged in serial flow order, thecompressor section and turbine section together defining a core airflowpath; a rotary component rotatable with at least a portion of thecompressor section and with at least a portion of the turbine section;an electric machine coupled to the rotary component at least partiallyinward of the core air flowpath along the radial direction; and a cavitywall defining at least in part a buffer cavity, the buffer cavitysurrounding at least a portion of the electric machine to thermallyinsulate the electric machine.
 2. The gas turbine engine of claim 1,wherein the buffer cavity extends from a location forward of theelectric machine to a location aft of the electric machine.
 3. The gasturbine engine of claim 1, wherein the buffer cavity is positioned atleast partially between the electric machine and the core air flowpath.4. The gas turbine engine of claim 1, wherein the electric machine ismounted at least partially within or aft of the turbine section alongthe axial direction.
 5. The gas turbine engine of claim 1, furthercomprising: a cooling duct in airflow communication with the buffercavity for providing a cooling airflow to the buffer cavity.
 6. The gasturbine engine of claim 5, further comprising: a fan positioned at aforward end of the gas turbine engine, wherein the cooling duct isadditionally in airflow communication with at least one of the fan orthe compressor section.
 7. The gas turbine engine of claim 1, furthercomprising: an electric machine lubrication system, wherein the cavitywall further defines at least in part an electric machine sump, whereinthe electric machine lubrication system is in fluid communication withthe electric machine sump for providing a thermal fluid to the electricmachine sump.
 8. The gas turbine engine of claim 7, wherein the electricmachine sump is positioned opposite the cavity wall of the buffercavity.
 9. The gas turbine engine of claim 7, further comprising: anengine lubrication system, wherein the engine lubrication systemoperates independently of the electric machine lubrication system. 10.The gas turbine engine of claim 1, wherein the electric machine is anelectric generator.
 11. The gas turbine engine of claim 10, wherein theelectric generator is a permanent magnet electric generator comprising aplurality of permanent magnets.
 12. The gas turbine engine of claim 11,further comprising: a cooling system operable with the buffer cavity,wherein the plurality of permanent magnets each define a Curietemperature limit, and wherein the cooling system maintains atemperature of each of the permanent magnets below the Curie temperaturelimit.
 13. The gas turbine engine of claim 12, wherein the coolingsystem maintains a temperature of the permanent magnets below at leastabout a 50° F. limit of the Curie temperature limit.
 14. The gas turbineengine of claim 1, wherein the electric machine is mounted coaxiallywith the rotary component.
 15. The gas turbine engine of claim 1,wherein the cavity wall includes insulation.
 16. The gas turbine engineof claim 1, wherein the buffer cavity substantially completely surroundsthe electric machine.
 17. A propulsion system for an aeronautical devicecomprising: an electric propulsor; and a gas turbine engine defining aradial direction and an axial direction, the gas turbine enginecomprising a compressor section and a turbine section arranged in serialflow order, the compressor section and turbine section together defininga core air flowpath; a rotary component rotatable with at least aportion of the compressor section and with at least a portion of theturbine section; an electric machine coupled to the rotary component atleast partially inward of the core air flowpath along the radialdirection, the electric machine electrically connected to the electricpropulsor; and a cavity wall defining at least in part a buffer cavity,the buffer cavity surrounding at least a portion of the electric machineto thermally insulate the electric machine.
 18. The propulsion system ofclaim 17, wherein the buffer cavity extends from a location forward ofthe electric machine to a location aft of the electric machine.
 19. Thepropulsion system of claim 17, wherein the buffer cavity is positionedat least partially between the electric machine and the core airflowpath.
 20. The propulsion system of claim 17, further comprising: anelectric machine lubrication system, wherein the cavity wall furtherdefines at least in part an electric machine sump, wherein the electricmachine lubrication system is in fluid communication with the electricmachine sump for providing a thermal fluid to the electric machine sump.